U.S. patent number 5,483,894 [Application Number 08/364,905] was granted by the patent office on 1996-01-16 for integral missile antenna-fuselage assembly.
This patent grant is currently assigned to Hughes Missile Systems Company. Invention is credited to Andrew B. Facciano, Ronald N. Hopkins, Rodney H. Krebs, James L. Neumann, Oscar K. Ohanian.
United States Patent |
5,483,894 |
Facciano , et al. |
January 16, 1996 |
Integral missile antenna-fuselage assembly
Abstract
An integral missile antenna-fuselage assembly (50) is provided
for integration into an armament missile (12) which carries primary
missile loads, houses internal electronic assemblies, provides
mounting surface zones for external sensor antennas (71) , and
protects sensitive antenna components from supersonic aerodynamic
heating. Each end of the fuselage assembly (50) is formed from a
fastener ring (52,54) having a circumferential recess (84,86) which
receives a filament wound main structure (60) to form the missile
fuselage tube. Preferably, a titanium liner (58) is first joined to
each fastener ring with a step-lap joint (94,96) along which it is
adhesively bonded. The liner (58) and adjacent fastener ring
portions (52,57) provide a mandrel on which a graphite/Bismaleimide
(BMI) resin pre-preg is filament wound and co-cured to form the
integral fuselage (60). A plurality of Graphite/BMI doublers
(62,63,64,65) are axisymmetrically positioned on the fuselage
external surface to form four antenna cavities (66,67,68,69) which
receive antennas (71) therein. Subsequently, antenna spacers
(72,73,74,75) encase the antennas (71) about which a radome
overwrap (70) is filament wound with a Quartz/BMI pre-preg. The
entire structure (70) is then integrally cured to the internal
fuselage (60) and antenna spacers (72,73,74,75) afterwhich it is
surface treated (76) and overcoated (78).
Inventors: |
Facciano; Andrew B. (Tucson,
AZ), Hopkins; Ronald N. (Superior Township, AZ), Krebs;
Rodney H. (Tucson, AZ), Neumann; James L. (Tucson,
AZ), Ohanian; Oscar K. (Tucson, AZ) |
Assignee: |
Hughes Missile Systems Company
(Los Angeles, CA)
|
Family
ID: |
23436612 |
Appl.
No.: |
08/364,905 |
Filed: |
December 27, 1994 |
Current U.S.
Class: |
102/293;
102/377 |
Current CPC
Class: |
F42B
15/00 (20130101) |
Current International
Class: |
F42B
15/00 (20060101); F42B 015/36 () |
Field of
Search: |
;102/293,377,473,374 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Jordan; Charles T.
Assistant Examiner: Wesson; Theresa M.
Attorney, Agent or Firm: Brown; Charles D. Heald; Randall M.
Denson-Low; Wanda K.
Claims
What is claimed is:
1. An assembly for use in an armament missile constructed from a
plurality of joined-together sections, said assembly
comprising:
a missile fuselage tube constructed of a composite material having
reinforcing fibers impregnated with resin;
a fastener ring having an outer rim portion with a radially inward
extending circumferential recess formed therein for receiving at
least ends of the fibers;
circumferential means surrounding the ends of the fibers to secure
the ends of the fibers within said rim portion recess; and
said resin further impregnating the ends of the fibers and the
circumferential means to bond the tube to the ring.
2. The structural joint of claim 1 wherein said resin comprises
Bismaleimide (BMI) resin.
3. The structural joint of claim 2 wherein said reinforcing fibers
and said circumferential means comprise graphite fibers.
4. An aft fuselage assembly for use in constructing a
multiple-section armament missile, the assembly comprising:
a first fastener ring having an outer rim portion with a radially
inward extending circumferential recess formed therein;
a second fastener ring having an outer rim portion with a radially
inward extending circumferential recess formed therein;
a liner extending between said first and second fastener rings
which retains said first and second rings in spaced-apart relation,
said liner affixed to said first fastener ring at a first end and
said second fastener ring at a second end; and
a filament wound main structure provided by at least one nested
enforcing fiber received on said liner and radially inwardly
received in each of said rim portion recesses, said fiber
thereafter wetted-out with resin to form a cured resin matrix
laminate structure which is recess trapped on said first and second
fastener rings at either end.
5. The assembly of claim 4 further comprising at least one doubler
received on an exterior surface of said main structure, at least
one antenna spacer which is constructed and arranged to provide at
least one axisymmetric antenna cavity therein and which cooperates
with said doubler to define a circumferential outer surface, and a
quartz overwrap further provided thereabout, wherein said overwrap
is wetted-out with resin and heated to form a cured resin matrix
laminate structure.
6. The assembly of claim 5 wherein said doubler comprises a
pressure-cured graphite composite.
7. The assembly of claim 5 wherein said overwrap comprises a
filament wound quartz pre-impregnated Bismaleimide (BMI) resin
composite.
8. The assembly of claim 5 further comprising a connector
through-hole provided in one of said fastener rings communicating
between a liner interior and said antenna cavity, when assembled,
and providing a passage for passing antenna cables
therethrough.
9. The assembly of claim 4 wherein one of said fastener rings
comprises a metal ring with an outer rim portion having a radially
inward extending circumferential recess formed therein and a resin
transfer molded composite insert assembly in-place molded to said
metal ring within said circumferential recess of said metal ring so
as to be recess trapped for rigid attachment therebetween.
10. The assembly of claim 8 further comprising an umbilical cavity
provided in one of said fastener rings and said main structure for
communicating between said liner interior and an exterior of the
aft fuselage assembly, wherein provision is made for
through-passage of antenna cables in a harness umbilical retained
on an armament missile exterior.
11. The assembly of claim 4 wherein at least one of said fastener
rings comprises a metal bolt ring having a radially inwardly
extending circumferential outer groove and a separate
circumferential composite rim structure which is affixed to said
bolt ring by forming said rim in entrapped engagement with said
bolt ring outer groove, wherein said rim portion recess is provided
in said composite rim structure.
12. The assembly of claim 4 wherein said liner comprises a metal
tube.
13. The assembly of claim 12 wherein said metal tube comprises
steel.
14. The assembly of claim 12 wherein said liner comprises titanium
which provides EMI and gas permeability shielding, and electrical
ground continuity therealong.
15. The assembly of claim 12 wherein said liner comprises a metal
tube which provides leakage prevention and EMI shielding, and metal
foil is co-cured on an internal surface of said composite rim
structure for providing further EMI and gas permeability shielding,
and electrical ground continuity throughout the aft fuselage
assembly.
16. The assembly of claim 15 wherein said liner comprises
titanium.
17. An armament missile constructed from a plurality of assembled
components comprising:
a first missile section;
a second missile section;
a third missile section disposed between said first and second
missile sections comprising a first fastener ring having an outer
rim portion with a radially inward extending circumferential recess
formed therein,
a second fastener ring having an outer rim portion with a radially
inward extending circumferential recess formed therein,
a liner extending between said first and second fastener rings
which retains said first and second rings in spaced-apart relation,
said liner affixed to said first fastener ring at a first end and
said second fastener ring at a second end, and a filament wound
main structure provided by at least one nested enforcing fiber
received on said liner and radially inwardly received in said rim
portion recesses, said fiber thereafter wetted-out with resin to
form a cured resin matrix laminate structure which is recess
trapped on said first and second fastener rings at either end.
18. The armament missile of claim 17 wherein one of said fastener
rings comprises a metal ring with an outer rim portion having a
radially inward extending circumferential recess formed therein and
a resin transfer molded composite insert assembly in-place molded
to said metal ring along said circumferential recess so as to be
recess trapped for rigid attachment therebetween.
19. The armament missile of claim 17 wherein a plurality of bolt
holes are provided in said second fastener ring for fixing said
third missile section to said adjoining second missile section.
Description
BACKGROUND OF THE INVENTION
1. Technical Field
This invention relates generally to a fuselage construction for an
armament missile and, more particularly, to an integral missile
antenna-fuselage assembly.
2. Discussion
Aft fuselage assemblies for use in constructing multiple section
armament missiles are known in the art which function doubly as a
primary structural member and a missile antenna housing. To this
end, armament missiles are generally constructed from a plurality
of joined-together sections. Each intermediate section includes a
pair of fastener joints provided one at each end of a cylindrical
section skin to form a missile section. Typically, an armament
missile from tip-to-tail has a guidance section, an armament
section, a propulsion section, and a control section. The aft end
of the guidance section is further sub-divided to include an aft
fuselage which joins the guidance section to the armament
section.
Accordingly, the aft fuselage section must carry primary vehicle
loads through the missile air frame in between the guidance section
and armament section. Likewise, the aft fuselage section must house
antenna components which form part of the guidance section to
control the missile in-flight.
It is therefore desirable to provide an improved aft fuselage for
the guidance section of an Advanced Medium Range Air-to-Air Missile
(AMRAAM), or guided missile which reduces cost and simplifies
manufacturing through part consolidation. In addition, it is
further desirable to eliminate a secondary process presently
utilized for incorporating antenna components onto a missile
surface. In particular, it is desirable to eliminate secondary
steps in incorporating an antenna in the fuselage, consolidating
common features from the fuselage, and integrating fabrication
steps which simplify the fuselage design and streamline its
production. It is further desirable to enhance product reliability
and repeatability. Other further desirable features include
improving material efficiency to obtain a greater air frame
capability as a missile structure and as an antenna radome.
SUMMARY OF THE INVENTION
In accordance with the teachings of the present invention, an
Integral Missile Antenna-Fuselage Assembly (IMAFA) is provided
which is designed to carry primary missile loads, house internal
electronic assemblies, provide mounting surface zones for external
sensor antennas, and protect sensitive antenna components from
supersonic aerodynamic heating. The antenna-fuselage assembly
includes a structural joint which joins together a pair of fastener
rings at opposite ends of a filament wound main structure to form a
missile fuselage tube. A titanium liner is preferably first joined
to each fastener ring with a scarf joint along which it is
adhesively bonded. The liner and an adjacent flange portion on each
fastener ring form a mandrel on which a Graphite/Bismaleimide (BMI)
resin pre-preg is filament wound and co-cured to form an integral
fuselage therebetween. A radially inwardly extending
circumferential recess provided on each fastener ring rim receives
a filament winding therein which traps the integral fuselage to
each fastener ring subsequent to curing. In a preferred embodiment,
the integral fuselage is co-cured with four uni-directional
Graphite/BMI doublers which are axisymmetrically positioned on the
external surface to form four Target Detection Device (TDD) antenna
cavities which receive antennas therein. Subsequently, four antenna
spacers enclose the antennas to form an external cylindrical
surface thereabout. Finally, a radome overwrap is filament wound
with Quartz/BMI pre-preg which is subsequently integrally cured to
the internal fuselage and antenna spacers and post cured prior to
surface treatment with polyurethane paint overcoat.
BRIEF DESCRIPTION OF THE DRAWINGS
Other objects and advantages of the present invention will become
apparent to those skilled in the art upon reading the following
detailed description and upon reference to the drawings in
which:
FIG. 1 is a perspective view of an AMRAAM, or guided missile with a
prior art aft fuselage dome assembled in the missile;
FIG. 2 is a vertical side view with portions shown in breakaway of
the prior art aft-fuselage as shown in FIG. 1 without the overwrap
and TDD antennas;
FIG. 3 is a partial sectional view of the prior art aft-fuselage
taken generally along 3--3 of FIG. 2 including the overwrap and TDD
antennas;
FIG. 4 is a partial centerline-sectional view of an integral
missile antenna-fuselage assembly in accordance with the preferred
embodiment of the present invention for use with the missile of
FIG. 1;
FIG. 5 is a somewhat diagrammatic sectional view depicting fiber
orientation in constructing the trapped taper joint on the aft
fastener ring structure of FIG. 4;
FIG. 6 is a partial vertical centerline-sectional view depicting an
alternative construction for joining the titanium inner liner to
the forward fastener ring than that already shown in FIG. 4;
FIG. 7 is a vertical centerline-sectional view of the aft fastener
ring including a Resin Transfer Molded (RTM) insert with an
integral umbilical cavity;
FIG. 8a is a cross-sectional view taken along line 8--8 of FIG. 2
depicting the prior aft-fuselage at the location of the electronics
unit assembly; and
FIG. 8b is a cross-sectional view corresponding with that shown in
FIG. 8a depicting the aft-fuselage of FIG. 4 in cross-section.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
An existing Guidance Section (GS) aft-fuselage 10 for the Advanced
Medium Range Air-to-Air Missile (AMRAAM) 12 is provided in FIG. 1
in accordance with the prior art. The prior art aft fuselage 10 as
shown in FIG. 2 is constructed and assembled with three cylindrical
subcomponents 14, 16, 18 having doubler reinforcements 20, 22, 24
therealong. The first subcomponent is an aft fuselage skin 14
formed from a sheet of titanium which forms the walls of the
fuselage. A forward flange 16 is machined from bars of annealed
titanium to define a first end of the fuselage. Art aft housing 18
is formed from a titanium investment cast structure to define the
opposite end of the fuselage. Aft fuselage skin 14 is preferably
formed in two halves which are subsequently joined together to
define a cylinder having longitudinal surface cavities 25 stamped
therein for supporting Target Detection Device (TDD) antennas.
According to the prior art, the aft fuselage skin is formed in two
halves by a pair of mating skin sections 26 and 28 which are welded
together along their longitudinal seams. Furthermore, the forward
flange 16 and aft housing 18 are circumferentially electronbeam
welded to opposite ends of the fuselage skin. However, to ensure
weld integrity full radiographic and ultrasonic inspections must be
made of each weld, and the entire structure must be helium leak
tested.
Furthermore, the plurality of doublers 20, 22, 24 formed from
titanium sheet metal are spot welded to the fuselage skin 14
in-between the antenna cavities 25 for reinforcement purposes.
Accordingly, all the aforementioned welds must be heat treated to a
temperature of approximately 1,100.degree. F. for about 120 minutes
in order to relieve stresses in the welds.
Following welding and heat treating of the prior art AMRAAM aft
fuselage section 10, eight TDD antennas 30 with coax cable
connectors are installed into the skin cavities 25 with Kapton tape
32 manufactured by DuPont de Nemours, E. I., & Co., Inc. As
shown in FIG. 3, a QUARTZ/POLYIMIDE (Qz/PI) spacer 24 is then
positioned over the antennas using Kapton tape in order to
complementarily shape the fuselage skin into an external
cylindrical shape. As shown in FIG. 8a the fuselage and antenna
assembly is then wet wound with a Qz/PI overwrap 36. However, this
technique is very labor intensive, complex to process, and very
costly per unit section. Furthermore, internal pressurization
during helium leak testing has been difficult to maintain when
using electron beam and doubler spot welds during assembly. The
Qz/PI overwrap is not fully cured in practice since the TDD
antennas can become dimensionally unstable and fail when heated
over 500.degree. F. which prevents fully curing the overwrap.
Furthermore, internal voids and surface cracks frequently form
which necessitates the application of a 0.005 inch thick Epoxylite,
an epoxy and solids filler adhesive sold by Epoxylite Corporation,
9400 Toledo Way, Irvine, Calif. 92713-9671, overwrap sealant, to
seal the voids and surface cracks. However, the epoxylite overwrap
sealant decomposes and burns in the range of
500.degree.-600.degree. F. This temperature restriction further
prevents the full curing of the Qz/PI overwrap.
FIG. 1 illustrates the major sections of the AMRAAM 12 including
the prior art aft fuselage 10 positioned between a GS forward
fuselage 38 and an armament section 40. The GS forward fuselage
houses a Terminal Seeker and radar transmitter unit (not shown).
Correspondingly, the prior art GS aft-fuselage houses the
Electronic Unit (EU) Assembly, the Inertial Reference Unit (IRU)
and the TDD Electronics and Antennas (not shown). Bending loads
generated by the forward and aft GS assemblies are transmitted
through the GS aft-fuselage Missile Station (MS) "55", designated
by numeral 44. The maximum bending moment at MS "55" is 1,015
lbs-inch which occurs as a result of a launch adapter unit (LAU)-92
eject launch. The forward pylon and eject launcher captive carry
feature is provided by a forward hanger 46 and hook 48 located at
the aft end of the armament section. Accordingly, all forward
missile vibration loads which are generated from a captive carry
aerodynamic buffet are transmitted through the aft-fuselage
structure to the warhead hanger and hook assembly, namely, hanger
46 and hook 48. The GS aft-fuselage is designed to withstand
missile free flight, eject launch, and captive carry fatigue loads
and extreme Air-to-Air Missile (AAM) thermal environments with
sufficient structural margin to ensure operation reliability. In
addition, the GS aft-fuselage provides the EU Electromagnetic
Interference (EMI) shielding and atmospheric isolation, the TDD
antennas mounted on an external mounting surface, and thermal
insulation for enveloping all of the electronic assemblies. As a
result, the GS aft-fuselage is the most significant and complex
vehicle fuselage assembly on AMRAAM, and the most expensive to
fabricate.
Turning now to FIGS. 4 and 5, an Integral Missile Antenna Fuselage
Assembly (IMAFA) 50 is shown in accordance with the present
invention. IMAFA 50 is substituted for the prior art GS aft
fuselage 10 where it is assembled into the missile 12. The
antenna-fuselage assembly 50 is shown in cross-section in order to
illustrate the various components utilized in constructing the
assembly. A forward joint ring 52 and an aft joint ring-insert
assembly 54 are simultaneously bonded to a near
cylindrical-hydroformed titanium or corrosion resistant steel
(CRES) structural liner. The aft joint ring-insert assembly 54
provides a fastener ring and is formed from a titanium joint ring
56 and an Resin transfer Molded (RTM) insert assembly 57
constructed from a RTM structure. Preferably, rings 52 and 56 are
machined from titanium. A plurality of circumferentially spaced
apart bolt holes 59 (only one of which is shown) are provided in
each ring for fastening to respective adjoining missile sections.
Alternatively, each ring is machined from corrosion resistant
steel. Forward joint ring 52 is located at Missile Station (MS)
"32", identified as numeral 42 in FIG. 1, on the AMRAAM missile,
and aft joint ring assembly 54 is located in the vicinity of
missile station (MS) "55", numeral 44, of the AMRAAM missile. The
RTM composite insert assembly is fabricated preferably from a
graphite fabric preform, injected with a Bismaleimide (BMI) resin
which is integrally formed onto the aft joint ring assembly 54.
Preferably, a near cylindrical, hydroformed titanium liner 58 is
simultaneously bonded to both the forward joint ring 52 and aft
joint ring-insert assembly 54 with a structural adhesive. The liner
58 is preferably 0.015 to 0.020 inches thick and functions as a
built-in filament winding mandrel which minimizes the cost of
having to utilize a separate mandrel during construction of the aft
fuselage assembly 50. Furthermore, the liner 58 provides the
internal EU assembly with EMI and gas permeability shielding, and
forms an integral, isotropic compression layer for the primary
fuselage structure. Alternatively, the liner 58 can be formed from
corrosion resistant steel (CRES).
A filament wound internal fuselage main structure 60 is formed over
the liner 58 and portions of ring 52 and ring assembly 54. The
internal fuselage main structure 60 provides primary load carrying
structure for fuselage assembly 50, and is fabricated by filament
winding Graphite/BMI pre-preg onto the resulting mandrel assembly
formed by liner 58, ring 52 and ring assembly 54. Preferably, a
structural adhesive is applied to the mandrel assembly prior to
filament winding the pre-preg. The internal fuselage main structure
60 is then co-cured with four uni-directional Graphite/BMI doublers
62-65 which are axisymmetrically positioned on the external surface
formed by structure 60 which assists to define four TDD antenna
cavities 66-69 circumferentially spaced apart thereabout.
As shown in FIG. 8b, eight TDD antennas 71 are placed into the
cavities 66-69, with two antennas per cavity. Four QZ/BMI antenna
spacers 72-75 are added to enclose the antennas and form an
external cylindrical surface. A radome overwrap, or QZ/BMI overwrap
70, is filament wound about the antenna spacers and doublers using
a QZ/BMI pre-preg and integrally cured at 350.degree. F. to the
internal fuselage and antenna spacers, then post-cured at
475.degree. F. to finish the IMAFA 50 prior to surface treatment
and application of a polyurethane overcoat 78.
An innovative structural feature on the fuselage assembly 50 is the
use of a trapped fiber, taper joint design 80, 82 at the aft and
forward interfaces between the main structure 60 where it engages
with the ring 52 and ring-insert assembly 54, respectively, as
exhibited in FIG. 4. FIG. 5 schematically illustrates construction
of a structural interface, namely fiber trap joint 82 formed on
ring insert assembly 54 FIG. 5 schematically depicts fiber trap
joint 82 which is formed in forward aft joint ring insert assembly
54. The internal fuselage main structure 60 is circumferentially
hoop wound about the liner 58, and further wound into a fiber trap
90, comprising a radially inwardly extending circumferential
recess. Alternatively, structure 60 can be formed from a cloth
weave such as a fiberglass cloth, or graphite cloth. Preferably, at
least one circumferential fiber 92 is subsequently
circumferentially wound over the filament windings to trap them
into the fiber trap 90 prior to wet-out or impregnation with a
resin in which it is cured.
In order to facilitate winding of main structure 60, liner 58 is
first adhesively retained to the forward joint ring 52 and the aft
joint ring-insert assembly 54 at either end. A step-lap joint 94 is
formed in joint ring 52 for receiving one end of the liner. A
second step-lap joint 96 is formed in RTM insert 57 for receiving
the opposite end of liner 58. Preferably, the liner is trapped and
bonded onto each joint ring 52 and 54 with structural adhesive to
form bond joint 84 and 86, respectively, in order to obtain
compressive strength therethrough.
The filament wound structure 60 is then wound onto the liner 58 and
inside the joint ring fiber trap joints 80 and 82 where further
filament windings form circumferential fibers 92 which trap
structure 60 therein. Alternatively, main structure 60 can be
formed from a fabric weave, such as fiberglass cloth which is
subsequently retained inside the fiber trap joints 80 and 82 with a
wrapping of circumferential fibers 92 about the cloth. The wound
structure 60 locks onto the rings 52 and 54 at fiber trap joints 80
and 82, respectively, to carry both compressive and tensile
loads.
Preferably, a heat-cured structural adhesive 98 is first applied to
all bond joint interfaces, namely, the joint between ring 56 and
RTM insert 57, between ring 52 and liner 58, and between insert 57
and liner 58, as well as in the fiber traps 90. As a result, the
primary composite structure adheres to the metallic liner and the
tapered joint interfaces which augments the compressive load
carrying capability of the liner. By combining the trap fiber,
taper joint design with the liner step-lap joint, a more
conservative configuration is provided for joining a main fuselage
structure 60 to a joint ring 52 and a joint ring assembly 54.
Therefore, an adequate design margin of safety is ensured which
meets the severe eject launch and captive carry fatigue
environments normally encountered with such a missile.
FIG. 6 depicts an alternative construction for the forward joint on
IMAFA 50. A modified forward joint ring 52' has a modified step-lap
joint 94' which is adhesively bonded to a modified titanium liner
58. An internal fuselage main structure 60' is filament wound about
the liner and joint ring, including a fiber trap joint 80' to bond
the main structure 60' to the forward joint ring 52'. Subsequently,
doublers 62, identical to those used in the preferred joint
construction, are received over a main structure 60' afterwhich
overwrap 70 is received and cured.
FIG. 7 depicts a selected cross section of the ring/insert assembly
54, including Graphite/BMI resin transfer molded insert 57. An
umbilical cavity 100 and a fill drain port 102 formed in insert 57
are shown in cross section. The umbilical cavity 100 allows
connection of an electronic unit (EU) motherboard housed within the
fuselage assembly 50 with a missile harness umbilical assembly 104
affixed to the missile exterior. As shown in FIG. 1, the umbilical
assembly 104 extends from the missile GS 37, namely the rear
portion of the aft fuselage 50, to the missile control section 41.
Additional umbilical cavities (not shown) are provided on the
armament section 40, propulsion section 39, and control section 41
for wiring to the umbilical assembly 104.
As shown in FIG. 7, the RTM insert 57 is thicker than the
Graphite/BMI filament wound skin 60 which compensates for
structural discontinuities normally encountered at a structural
joint to provide a stiff, extremely stable Inertial Reference Unit
(IRU) platform to MS "55", numeral 44. Numerous bosses, material
standoffs, connector through holes, and fastener inserts are
incorporated on the internal surface to mount the IRU, TDD
Electronics and Coax Cable Assemblies inside the aft fuselage
50.
A metallic foil 106 is preferably co-cured on internal surface of
RTM insert 57 to provide EMI and gas permeability shielding , and
electrical ground continuity throughout the length of the aft
fuselage 50. Perforations are provided in the foil 106 for through
passage of bosses and access to umbilical cavities and sockets.
Alternatively, surface sealants and electrically conductive paints
can be substituted for foil 106.
The aft joint ring/insert assembly 54 is joined together with a
mechanical locking joint which augments structural adhesive applied
to the joined surfaces. A circumferential groove 108 is provided in
the joint ring 56 into which the RTM insert is molded which traps
the ring and insert together. Furthermore, groove 108 terminates in
the region of the umbilical cavity 100 and a local groove 110
couples the ring and insert together in the region of the cavity
100. The mechanical joint formed therebetween functions
mechanically similarly to the trapped fiber, taper fuselage joints
80 and 82. In each of these joints, catastrophic failure will only
occur after the mechanically superior graphite fibers are fractured
and break, instead of relying solely on the adhesive shear strength
of a bonded joint configuration.
The IMAFA composite design for aft fuselage 50 avoids material
stress concentrations and load path discontinuities associated with
traditional fasteners. An attempt is made to incorporate uniform
stress path characteristics in critical structural interfaces with
composite material in order to eliminate any weak-link in an
aerospace structure. Therefore, bond joints 84 and 86 at Missile
Stations "32" and "55" have thin flanges, closely spaced
countersunk holes 59 fully stressed in bearing and shear, and
flathead screws torqued to the maximum allowable levels.
Countersunk holes are position toleranced very tight to minimize
stress concentration induced fatigue failures. Missile Stations
"32" and "55" are also exposed to severe flight temperatures and a
wide range of corrosive elements resulting from airborne captive
carry. The forward and aft fiber trap joints 80 and 82 conflict
with the design guidelines established within the industry for
composite fastener applications. Therefore, aft fuselage 50
additionally incorporates the titanium, or CRES, ring structures 52
and 54 at Missile Stations "32" and "55" to meet the guidelines, as
well as to form a mandrel on which structure 60 is formed.
The design of aft fuselage 50 is optimized to enhance structural
reliability and material efficiency. Fuselage 50 has features
designed to perform multiple roles or provide secondary features
which augment their primary features. In use, fuselage 50 is
completely sealed with adjacent missile sections and various
connectors and fasteners, for example bolt holes 59, are sealed
with a polysulfide sealant. The sealed fuselage, which houses
missile electronics, is then pressurized with nitrogen to provide a
zero humidity environment for the high power microwave electronics.
In combination with the built-in shielding, the electronics are
protected from both humidity and magnetic fields created by corona
effects about the missile. FIG. 8-b depicts aft fuselage 50 in
cross-sectional view at the location of the electronic unit (not
shown). Likewise, the prior art aft fuselage 10 is also shown in
FIG. 8-a at the same location. Doublers 62-65 and antenna cavities
66-69 are clearly visible in FIG. 8-b. The thickness and filament
ply angles for the internal fuselage main structure 60 are
preferably determined by structural Finite Element Model (FEM)
analysis, preferably to match the natural vibration frequencies and
mode shapes of the current GS aft-fuselage 10. Preliminary analysis
has shown that a preferred composite laminate thickness and ply
angle to be approximately 0.050 inches and .+-.20 degrees,
respectively. The doublers are positioned between the internal
fuselage 60 and Qz/BMI overwrap 70 to provide fuselage stiffness
during eject launch, antenna cavity depth, and insulation for the
internal fuselage 60 from missile flight and captive carry thermal
transients. Radome overwrap 70 is integrally cured to the doublers
and antenna spacers to encapsulate the TDD antennas from
atmospheric humidity and form a cylindrical sandwich structure for
maximum load carrying capability. The radome overwrap 70 will
augment the bending inertia of the internal fuselage 60 to minimize
moment induced stresses during captive carry buffet and maximize
fatigue life.
Previous composite missile airframe fabrication experience lead to
the selection of BMI resin in constructing aft fuselage 50. Initial
work with glass reinforced BMI showed high temperature capability
and low cost. Hexcel F650 BMI resin presently appears to show the
best high temperature capabilities. Alternatively, Hexcel F655
toughened BMI and YLA RS-3 Poly Cyanate were found to be acceptable
resins for the internal fuselage 60 which improve damage tolerance
and fatigue durability.
It is to be understood that the invention is not limited to the
exact construction illustrated and described above, but that
various changes and modifications may be made without departing
from the spirit and scope of the invention as defined in the
following claims.
Thus, while this invention has been disclosed herein in combination
with particular examples thereof, no limitation is intended thereby
except as defined in the following claims. This is because a
skilled practitioner recognizes that other applications can be made
without departing from the spirit of this invention after studying
the specification and drawings.
* * * * *